1. Field of the Invention
The invention concerns a geocentric pointing three-axis stabilized satellite configuration for any orbit (possibly heliosynchronous, preferably but not necessarily steeply inclined and/or of low altitude); it is more particularly directed to the location and the kinetics of its solar generator.
2. Description of the Prior Art
A satellite has two parts: a platform including the various equipment necessary to the control and the operation of the satellite (including attitude and orbit control and electrical power supply systems) and a payload embodying various equipment for carrying out the specific mission of the satellite (these can include observation (remote sensing) or telecommunications equipment).
Satellite payloads require an input of electrical power usually supplied by a solar generator including photo-voltaic cells illuminated by the Sun (except when the satellite is in an eclipse).
The orientation of the solar generator(s) can be steered so that the solar cells are perpendicular to the Sun (or as close as perpendicular thereto as possible) in order to recover the maximal power and consequently to minimize the surface area of the solar generator required. The orientation depends on which side of the satellite carries the solar generator and its mechanism: it therefore depends on the configuration of the satellite.
In low orbit the increased diameter of the Earth at the equator and the resulting terms of the terrestrial attraction potential cause rotation--referred to as nodal regression--of the line of the nodes of the orbit (where the orbital plane intersects the equatorial plane) whose speed is given by an equation linking the semi-major axis a, the inclination i and (second order effects) the eccentricity e of the orbit.
In a heliosynchronous orbit nodal regression is used to exactly compensate the rotation of the Earth around the Sun (once per year) so that the satellite illumination conditions remain constant. This heliosynchronous property is particularly beneficial for optical observation satellites.
This kind of orbit considerably simplifies the problems of orienting the solar generator: a single rotation axis is sufficient to maintain the cells perpendicular to the Sun at all times.
However, although this type of orbit is well suited to optical observation missions, it can have undoubted drawbacks for other missions, due, in particular, to the fact that the satellite always overflies a given location at the same local time.
For this reason, in certain missions, it can be much better from the operational point of view to choose a non-heliosynchronous orbit (orbit of any kind); however, it can be shown that it is then always necessary to provide two degrees of freedom in rotation to keep the solar generator perpendicular to the Sun at all times. This is because it is necessary to compensate for the following two rotations:
nodal regression of the orbit relative to the direction of the Sun (non-heliosynchronism); and PA1 rotation of the satellite in its orbit so that one side is always pointed towards the Earth. PA1 the rotation axis of the single mechanism is at least approximately in the roll-yaw plane; this has the advantage of a configuration which is symmetrical about the roll-yaw plane, minimizing any disturbing torque due to atmospheric drag (in Earth orbit), solar radiation pressure and the gravity gradient, for example; PA1 the center of mass of the solar generator is at least approximately in the roll-yaw plane; this has the advantage of minimizing disturbing torques during rotation of the solar generator about Z; PA1 the rotation axis of the mechanism is near an edge of the anti-heavenly body side of the satellite body, which facilitates implementation of the wing and its arm in the stowed configuration; PA1 the wing is formed by an odd number of panels including a center panel to which the arm is connected, which facilitates stowing and deploying the array; PA1 the wing is formed of panels whose shape and dimensions are close to the shape and dimensions of the anti-heavenly body side, which facilitates stowing the combination of the satellite body and its folded solar generator under the nose-cone of the launch vehicle; PA1 the constant inclination is between 30.degree. and 35.degree. for low orbit altitudes between approximately 600 km and approximately 1,000 km, independently of the inclination of the orbit; PA1 the satellite includes an attitude control device including a non-null mean kinetic moment momentum wheel whose axis is at least approximately parallel to the pitch axis; this is possible because the configuration of the satellite of the invention as defined hereinabove minimizes disturbing torques about the roll and yaw axes; the pitch axis is virtually an inertial axis (bear in mind that the pitch axis must be kept perpendicular to the plane of the orbit) and any disturbing torques about the pitch axis can be compensated by the momentum wheel; the latter also controls attitude in pitch and, by virtue of its gyroscopic stiffness, coupled control of the roll and yaw angles; this has the advantage over prior art solutions of minimizing the number of wheels needed for satellite attitude control; and PA1 at least one heatsink or thermal radiator is disposed on the anti-heavenly body side or on the "heavenly body" side, benefiting from the large surface area that these sides can have; the anti-heavenly body side has a high heat rejection capacity, greater than that of the heavenly body side, but sees more of the Sun; equipment moderately sensitive to the thermal gradient can be installed on this side; components excessively sensitive to the thermal gradient are preferably installed on the heavenly body side.
The invention applies to this type of orbit, preferably but not necessarily a low orbit.
A satellite which rotates in its orbit is usually stabilized in a particular attitude chosen to suit its mission. In the case of geocentric satellite attitude, stabilization keeps a so called Earth side of the satellite facing towards the Earth, i.e. perpendicular to the geocentric direction (usually called the Z direction). The payload (antennas, instruments, optical devices, etc.) is usually located on the Earth side (or +Z side). The opposite side is called the anti-Earth side (or -Z side).
Depending on the geocentric stabilization mode adopted, the satellite may or may not retain a fixed attitude relative to the geocentric direction and the orbit. If the attitude relative to the orbit can vary, there is one degree of freedom in rotation about Z, which can be exploited to keep the solar generator perpendicular to the solar radiation. However, to fulfill their mission some payloads must remain in an appropriate attitude both relative to Z and relative to the orbit, which is incompatible with one degree of freedom in rotation about Z. The satellite must then be stabilized about three axes (not only the Z axis, but also the roll axis X and the pitch axis Y) relative to its orbit.
The invention applies to this type of satellite configuration.
Examples of satellites in low orbit are given in European Patent No. 0,195,553. Another is the METEOR satellite.
European Patent No 0,195,553 describes an orbital flight satellite with two possible modes of attitude pointing, namely an Earth-pointing (geocentric) mode and a Sun-pointing mode. The satellite has an elongate cylindrical body and two solar generator wings or arrays extending transversely to the body and articulated to a middle portion thereof. In Earth-pointing mode the axis of the body is parallel to the geocentric direction. The two wings are usually aligned with each other but can be disposed one beside the other if two bodies are fixed together.
To allow this the wings can rotate 90.degree. about the axis. In normal service, for tracking the solar direction, each wing also turns about its longitudinal axis and about an axis perpendicular to the plane of the axis of the body and the longitudinal axis. It is clear that any such solution is complex (there are three degrees of freedom for the arrays with necessarily complex drive devices, likely to lead to serious malfunctions); moreover, there is no provision to minimize shading of the wings by the body of the satellite or its equipment and the existence of any such risk of shading implies over-rating of the solar generator, and thus of the number of solar cells, to guarantee an adequate supply of electrical power to the remainder of the satellite at all times. Finally, this configuration implies the provision in the satellite body itself of a non-negligible space to allow for the various angular movements of the wings. This can lead to problems in the implementation of the payload.
The METEOR satellite has a satellite body and two solar generator wings aligned with each other transversely to the body and mounted on a gantry joined to the body by a guide mechanism mounted on the anti-Earth side and having a rotation axis coincident with the geocentric direction. The wings are inclined to the plane defined by the geocentric direction and the direction in which the wings extend. This configuration has drawbacks including the fact that, even though in service the wings require fewer degrees of freedom, their deployment (in combination with the gantry) necessitates complex movements, all the more so in that the body of the satellite is a cylinder which is elongate in the geocentric direction. What is more, the fact that the arrays and their gantry turn around the satellite body requires the body to have similar dimensions in orthogonal positions transverse to the geocentric direction.